The invention relates generally to electrical power and cooling systems for aircraft and more particularly to enabling high energy system operation using an integrated power and cooling unit for high performance aircraft.
Modern aircraft integrate a number of systems to perform the functions required for flight and operation. A propulsion engine provides power to the aircraft when in flight and also drives the main generator to provide electrical power, either during flight or when on the ground. In order to provide the emergency backup power in the event of main engine or main generator failures, aircraft have evolved to include a supplemental non-propulsion engine such as the auxiliary power unit (APU). Since the cooling system is a major function of an aircraft, it has been integrated with the APU to provide not only cooling but also power in the event that it is required if the main generator system failed.
This integrated system (APU and cooling system) is often referred to as an integrated power and cooling unit (IPCU) that not only provides pneumatic or electrical power for starting the propulsion engine; it also generates electrical power and provides conditioned cooling air to the aircraft, both during flight and while on the ground. With the increasing use of systems having high energy requirements on aircraft, the IPCU can also be used to help meet short term high peak power when needed so as to minimize over-sizing the main generator system.
FIG. 1 is a schematic block diagram of portions of an aircraft electrical power and cooling system. As shown in FIG. 1, a prior art IPCU 100 includes cooling turbine 102, compressor 104, starter/generator 106 and power turbine 108, all connected to a common drive shaft 110. In order for the IPCU to come up to operating speed, initial startup of IPCU 100 is driven by starter/generator unit 106 in starter mode using battery or ground power. After compressor 104 is capable of providing air for combustion then power turbine 108 generates enough power to sustain the power requirements with the system configured to burn fuel using combustor 112 and thus, continues to drive the power turbine to generate power. After IPCU 100 is started in combustion mode, it can provide electrical power for various aircraft systems through integrated control unit (ICU) 114 which sends the power to integrated power unit (IPU) Bus 115. When in normal flight mode, IPCU 100 is then configured to provide cooling air by running off of propulsion engine bleed air instead of power input from burning fuel. This is accomplished by means of cooling turbine 102, which also provides low pressure cool air to an avionics cooling system. This system includes heat exchangers 116 and 118 as well as pump 120. The avionics cooling system is used to provide temperature controlled air flow to the avionics equipment and cabin of the manned aircraft, as well as for other needs as understood by one of ordinary skill in the art.
Also in FIG. 1, main engine 122 is shown, together with an engine high pressure spool coupled starter/generator 124 and an engine low pressure spool driven generator 126. High pressure spool starter generator 124 is connected to engine 122 as well as to an electrical power distribution bus 128 through an inverter control unit 130. Low pressure spool generator (LP GEN) 126 is also connected to engine 122 as well as to an electrical power bus 132 through generator control unit (GCU) 134.
There are several connections between IPCU 100 and engine 122. High pressure, warm air from compressor 104 can be directed into fan duct heat exchanger 136 of engine 122 when operating in the cooling air mode. This air can also be directed into combustor 112 to generate more power by burning fuel and driving the power turbine 108 using valve 148. Compressor 104 also receives ambient air through input 142 when operating in ground operating mode or in-flight emergency mode.
FIG. 1 shows a closed loop system, which includes heat exchanger 138 that provides pre-cooled engine bleed air to compressor 104. The air is compressed by the compressor 104 and then cooled by the engine fan air through the fan duct heat exchanger 136. An additional heat exchanger 140 cools the air provided to the cooling turbine using the relatively cool air returned from the avionics heat exchanger 116. The air is expanded by cooling turbine 102 to generate very cold air to cool the avionics liquid cooling loop through avionics cooler 116. The cooling capacity of the system is determined by the exit air temperature and the mass flow rate of the cooling turbine. A closed loop system has the advantage of allowing lower bleed air usage which conserves fuel, however, IPCU system pressure is limited by the compressor 104 pressure ratio capability. In other words, in order for the air flow exiting from cooling turbine 102 to feedback to the compressor 104 through heat exchanger 140, the return pressure has to be higher that the replenish flow from the engine. This pressure is set by the operating parameters of compressor 104 when operating at a high altitude. Often, selector/regulator valves 144 and 146 are used to select the engine bleed air according to the operating altitude. Due to the high pressure ratio of the modern engine compressor, this limits the cooling turbine 102 discharging pressure and the temperature of the cooling flow.
In contrast, FIG. 2 depicts a prior art open loop system where heat exchanger 140 does not provide an input to compressor 104 but is controlled by a venting valve 150 then vented to the ambient condition. This open loop system allows a lower cooling temperature exit from the cooling turbine, which means a higher expansion ratio is available at high altitude. However, bleed air usage is higher and is limited by the flow and temperature required to cool the avionics. Traditionally, in order to regulate the flow and the cooling capability, the system exit flow pressure and thus the system operating speed is controlled by placing back pressure to the cooling turbine using an exhaust control valve 150. However, in order to deep discharge the cooling turbine 102, power must be absorbed by compressor or the starter/generator 106 on the same shaft of IPCU 100. Since this prior art system is not designed to use the starter/generator in generating mode to absorb the power, it is thus incapable of operating economically at a wide range of power and cooling capacities. Specifically, if there is peak power equipment that only requires peak power occasionally during the flight mission, then the system must be over sized to be capable of providing the maximum cooling capability. Therefore, the system is operating at a much lower setting and lower efficiency most of the time thus resulting in a less efficient system. In prior art systems, starter/generator 106 is only used during system startup, and not during cooling modes of the system.
Thus, a need exists for an improved integrated power and cooling system that is capable of providing efficient peak power and cooling not limited by the operating pressure of the closed loop system or the by the power balance required to maintain IPCU 100 main shaft speed using only the air control valves of the open-loop system.